U.S. patent application number 16/751298 was filed with the patent office on 2020-09-03 for conformal seal and vane bow wave cooling.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Sean D. Bradshaw, Shawn M. McMahon, Dennis M. Moura, Steven D. Porter, Noah Wadsworth, Christopher Whitfield, Kevin Zacchera.
Application Number | 20200277867 16/751298 |
Document ID | / |
Family ID | 1000004842867 |
Filed Date | 2020-09-03 |
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United States Patent
Application |
20200277867 |
Kind Code |
A1 |
Porter; Steven D. ; et
al. |
September 3, 2020 |
CONFORMAL SEAL AND VANE BOW WAVE COOLING
Abstract
A gas turbine engine includes a combustor. A turbine section is
in fluid communication with the combustor. The turbine section
includes a first vane stage aft of the combustor. A seal assembly
is disposed between the combustor and the first vane stage. The
seal assembly includes a first plurality of openings and the first
vane stage includes a second plurality of openings communicating
cooling airflow into a gap between an aft end of the combustor and
the first vane stage. A first vane stage assembly and a method are
also disclosed.
Inventors: |
Porter; Steven D.;
(Wethersfield, CT) ; McMahon; Shawn M.; (West
Hartford, CT) ; Zacchera; Kevin; (Glastonbury,
CT) ; Wadsworth; Noah; (Sturbridge, MA) ;
Whitfield; Christopher; (Manchester, CT) ; Bradshaw;
Sean D.; (White Plains, NY) ; Moura; Dennis M.;
(South Windsor, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
1000004842867 |
Appl. No.: |
16/751298 |
Filed: |
January 24, 2020 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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15690615 |
Aug 30, 2017 |
10584601 |
|
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16751298 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2260/20 20130101;
F01D 9/023 20130101; F01D 25/12 20130101; F01D 11/005 20130101;
F05D 2240/55 20130101; F01D 9/041 20130101; F05D 2220/32 20130101;
F05D 2240/11 20130101; F05D 2240/121 20130101 |
International
Class: |
F01D 9/02 20060101
F01D009/02; F01D 11/00 20060101 F01D011/00; F01D 25/12 20060101
F01D025/12; F01D 9/04 20060101 F01D009/04 |
Claims
1. A gas turbine engine comprising: a combustor; a turbine section
in fluid communication with the combustor, the turbine section
including a first vane stage defining a portion of a core flow path
aft of the combustor; a seal assembly disposed between the
combustor and a forward face of the first vane stage, the seal
assembly including a radially outer surface, a radially inner
surface, a plurality of circumferentially spaced apart slots and a
first plurality of openings that extend from the radially outer
surface to the radially inner surface at an angle relative to the
radially outer surface at each of the plurality of
circumferentially spaced apart slots for communicating cooling
airflow into a gap between the aft end of the combustor and the
forward face of the first vane stage; and a second plurality of
openings through the first vane stage communicating cooling airflow
into the core flow path.
2. The gas turbine engine as recited in claim 1, wherein the first
vane stage includes a plurality of vanes with each of the plurality
of vanes including a leading edge and the plurality of
circumferentially spaced apart slots are located adjacent the
leading edge of each of the plurality of vanes.
3. The gas turbine engine as recited in claim 2, wherein the first
plurality of openings extend through the seal assembly to
communicate cooling airflow into a first set of the plurality of
slots and the second plurality of openings extend through the first
vane stage to communicate cooling airflow into a second set of the
plurality of slots.
4. The gas turbine engine as recited in claim 3, wherein the first
plurality of slots and the second plurality of slots alternate
circumferentially about the circumference of the first vane
stage.
5. The gas turbine engine as recited in claim 3, wherein the first
plurality of slots and the second plurality of slots are the same
size such that cooling air is communicated to each of the first and
second plurality of slots from cooling holes in both the first vane
stage and the seal assembly.
6. The gas turbine engine as recited in claim 1, wherein the seal
assembly includes an aft face that seals against a forward face of
the first vane stage, the aft face including a wearing end portion
extending axially aft from the seal assembly and configured to wear
down during initial operation to provide a seal against the forward
face.
7. The gas turbine engine as recited in claim 1, wherein the seal
assembly includes an alignment slot that aligns the seal assembly
circumferentially with the first vane stage.
8. A first vane stage assembly for a gas turbine engine comprising:
a first vane stage including an axial face, the first vane stage
defining a portion of a core flow path; and a seal assembly
abutting the axial face and extending axially across a gap between
a combustor and the first vane stage, wherein the seal assembly
includes a first plurality of openings that extend from a radially
outer surface to a radially inner surface and the first vane stage
includes a second plurality of openings through the axial face for
communicating cooling airflow into the core flow path.
9. The first vane stage assembly as recited in claim 8, wherein the
seal assembly includes a plurality of slots disposed at spaced
apart circumferential positions corresponding with the leading edge
of vanes of the first turbine stage.
10. The first vane stage assembly as recited in claim 9, wherein
the first plurality of openings and the second plurality of
openings open into a corresponding one of the plurality of
slots.
11. The first vane stage assembly as recited in claim 10, wherein
the first plurality of openings and the second plurality of
openings are disposed in groups spaced apart circumferentially to
correspond with the circumferential positions of the plurality of
slots.
12. The first vane stage assembly as recited in claim 11, wherein
the first plurality of openings are in communication with a first
set of the plurality of slots and the second plurality of openings
are in communication with a second set of the plurality of slots
that is different than the first set of the plurality of slots.
13. The first vane stage assembly as recited in claim 8, wherein
the second plurality of openings extend at an angle through the
axial face of the first vane stage.
14. The combustor assembly as recited in claim 8, wherein the seal
assembly includes an alignment slot that aligns the seal assembly
circumferentially with the first vane stage.
15. A method of cooling an interface between a combustor and a
turbine vane stage comprising: assembling a seal across a gap
between a combustor and a turbine vane stage aft of the combustor;
and communicating cooling air flow into the gap through a first
plurality of openings in the seal and into a core flow path with a
second plurality of openings in the turbine vane stage.
16. The method as recited in claim 15, including forming the seal
to include a plurality of circumferential slots and aligning the
plurality of circumferential slots with a leading edge of turbine
vanes within the turbine vane stage.
17. The method as recited in claim 16, including grouping the first
plurality of openings and the second plurality of openings
circumferentially to correspond with the location of the plurality
of circumferential slots and the leading edge of the turbine vanes.
Description
CROSS REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. application Ser.
No. 15/690,615 filed on Aug. 30, 2017.
BACKGROUND
[0002] A gas turbine engine typically includes a fan section, a
compressor section, a combustor section and a turbine section. Air
entering the compressor section is compressed and delivered into
the combustion section where it is mixed with fuel and ignited to
generate a high-energy exhaust gas flow. The high-energy exhaust
gas flow expands through the turbine section to drive the
compressor and the fan section. The compressor section typically
includes low and high pressure compressors, and the turbine section
includes low and high pressure turbines.
[0003] An interface between the combustor exit and the first vane
stage can experience elevated temperatures at localized areas near
a leading edge of each vane. The interface between the combustor
exit and the first vane stage includes a gap. Bow wave phenomena at
the leading edge of each vane in combination with the gap can
result in elevated temperatures within and near the gap at this
location.
[0004] Turbine engine manufacturers continue to seek improvements
to engine performance including improvements to thermal, transfer
and propulsive efficiencies.
SUMMARY
[0005] In a featured embodiment, a gas turbine engine includes a
combustor. A turbine section is in fluid communication with the
combustor. The turbine section includes a first vane stage aft of
the combustor. A seal assembly is disposed between the combustor
and the first vane stage. The seal assembly includes a first
plurality of openings and the first vane stage includes a second
plurality of openings communicating cooling airflow into a gap
between an aft end of the combustor and the first vane stage.
[0006] In another embodiment according to the previous embodiment,
the first vane stage includes a plurality of vanes with each of the
plurality of vanes including a leading edge and the seal assembly
includes a plurality of slots disposed at circumferential positions
corresponding with the leading edge of each of the plurality of
vanes.
[0007] In another embodiment according to any of the previous
embodiments, the first plurality of openings extend through the
seal assembly to communicate cooling airflow into a first set of
the plurality of slots and the second plurality of openings extend
through the first vane stage to communicate cooling airflow into a
second set of the plurality of slots.
[0008] In another embodiment according to any of the previous
embodiments, the first plurality of slots and the second plurality
of slots alternate circumferentially about the circumference of the
first vane stage.
[0009] In another embodiment according to any of the previous
embodiments, the first plurality of slots and the second plurality
of slots are the same such that cooling air is communicated to each
of the first and second plurality of slots from cooling holes in
both the first vane stage and the seal assembly.
[0010] In another embodiment according to any of the previous
embodiments, the seal assembly includes a radially outer surface
and the first plurality of openings are angled relative to the
radially outer surface.
[0011] In another embodiment according to any of the previous
embodiments, the seal assembly includes aft face that seals against
a forward rib of the first vane stage.
[0012] In another embodiment according to any of the previous
embodiments, the seal assembly includes an alignment slot that
aligns the seal assembly circumferentially with the first vane
stage.
[0013] In another featured embodiment, a first vane stage assembly
for a gas turbine engine includes a first vane stage including an
axial face. A seal assembly abuts the axial face and extends
axially across a gap between a combustor and the first turbine vane
stage. The seal assembly includes a first plurality of openings and
the first vane stage includes a second plurality of openings
communicating cooling airflow into the gap.
[0014] In another embodiment according to any of the previous
embodiments, the seal assembly includes a plurality of slots
disposed at circumferential positions corresponding with the
leading edge of vanes of the first turbine stage.
[0015] In another embodiment according to any of the previous
embodiments, the first plurality of openings and the second
plurality of openings open into a corresponding one of the
plurality of slots.
[0016] In another embodiment according to any of the previous
embodiments, the first plurality of openings and the second
plurality of openings are disposed in groups spaced apart
circumferentially to correspond with the circumferential positions
of the plurality of slots.
[0017] In another embodiment according to any of the previous
embodiments, the first plurality of openings are in communication
with a first set of the plurality of slots and the second plurality
of openings are in communication with a second set of the plurality
of slots that is different than the first set of the plurality of
slots.
[0018] In another embodiment according to any of the previous
embodiments, the second plurality of openings extend at an angle
through the axial face of the first vane stage.
[0019] In another embodiment according to any of the previous
embodiments, the seal assembly includes an alignment slot that
aligns the seal assembly circumferentially with the first vane
stage.
[0020] In another featured embodiment, a method of cooling an
interface between a combustor and a turbine vane stage includes
assembling a seal across a gap between a combustor and a turbine
vane stage aft of the combustor. Cooling air flow is communicated
into the gap through a first plurality of openings in the seal and
a second plurality of openings in the turbine vane stage.
[0021] In another embodiment according to any of the previous
embodiments, includes forming the seal to include a plurality of
circumferential slots and aligning the plurality of circumferential
slots with a leading edge of turbine vanes within the turbine vane
stage.
[0022] In another embodiment according to any of the previous
embodiments, includes grouping the first plurality of openings and
the second plurality of openings circumferentially to correspond
with the location of the plurality of circumferential slots and the
leading edge of the turbine vanes.
[0023] Although the different examples have the specific components
shown in the illustrations, embodiments of this disclosure are not
limited to those particular combinations. It is possible to use
some of the components or features from one of the examples in
combination with features or components from another one of the
examples.
[0024] These and other features disclosed herein can be best
understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0025] FIG. 1 is a schematic view of an example gas turbine
engine.
[0026] FIG. 2 is a cross-sectional view of a portion of the gas
turbine engine.
[0027] FIG. 3 is a front view of a first turbine vane stage.
[0028] FIG. 4 is a perspective view of an interface between a
combustor and an example first turbine stage.
[0029] FIG. 5 is another perspective view of the interface between
the combustor and the example first turbine stage.
[0030] FIG. 6 is an axial front view of a conformal seal.
[0031] FIG. 7 is a front view of a portion of the example conformal
seal.
[0032] FIG. 8 is an enlarged view of a slot of the example
conformal seal.
[0033] FIG. 9A is a schematic illustration of an example cooling
air hole grouping.
[0034] FIG. 9B is another schematic illustration of another example
cooling air hole grouping.
DETAILED DESCRIPTION
[0035] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flow path B in a bypass duct defined within a
nacelle 18, while the compressor section 24 drives air along a core
flow path C for compression and communication into the combustor
section 26 then expansion through the turbine section 28. Although
depicted as a two-spool turbofan gas turbine engine in the
disclosed non-limiting embodiment, it should be understood that the
concepts described herein are not limited to use with two-spool
turbofans as the teachings may be applied to other types of turbine
engines including three-spool architectures.
[0036] The exemplary engine 20 generally includes a low speed spool
30 and a high speed spool 32 mounted for rotation about an engine
central longitudinal axis A relative to an engine static structure
36 via several bearing systems 38. It should be understood that
various bearing systems 38 at various locations may alternatively
or additionally be provided, and the location of bearing systems 38
may be varied as appropriate to the application.
[0037] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a first (or low) pressure compressor
44 and a first (or low) pressure turbine 46. The inner shaft 40 is
connected to the fan 42 through a speed change mechanism, which in
exemplary gas turbine engine 20 is illustrated as a geared
architecture 48 to drive the fan 42 at a lower speed than the low
speed spool 30. The high speed spool 32 includes an outer shaft 50
that interconnects a second (or high) pressure compressor 52 and a
second (or high) pressure turbine 54. A combustor 56 is arranged in
exemplary gas turbine 20 between the high pressure compressor 52
and the high pressure turbine 54. A mid-turbine frame 58 of the
engine static structure 36 is arranged generally between the high
pressure turbine 54 and the low pressure turbine 46. The
mid-turbine frame 58 further supports bearing systems 38 in the
turbine section 28. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via bearing systems 38 about the engine
central longitudinal axis A which is collinear with their
longitudinal axes.
[0038] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine
frame 58 includes airfoils 60 which are in the core airflow path C.
The turbines 46, 54 rotationally drive the respective low speed
spool 30 and high speed spool 32 in response to the expansion. It
will be appreciated that each of the positions of the fan section
22, compressor section 24, combustor section 26, turbine section
28, and fan drive gear system 48 may be varied. For example, gear
system 48 may be located aft of combustor section 26 or even aft of
turbine section 28, and fan section 22 may be positioned forward or
aft of the location of gear system 48.
[0039] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than about six (6), with an example embodiment being
greater than about ten (10), the geared architecture 48 is an
epicyclic gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3 and
the low pressure turbine 46 has a pressure ratio that is greater
than about five. In one disclosed embodiment, the engine 20 bypass
ratio is greater than about ten (10:1), the fan diameter is
significantly larger than that of the low pressure compressor 44,
and the low pressure turbine 46 has a pressure ratio that is
greater than about five 5:1. Low pressure turbine 46 pressure ratio
is pressure measured prior to inlet of low pressure turbine 46 as
related to the pressure at the outlet of the low pressure turbine
46 prior to an exhaust nozzle.
[0040] The geared architecture 48 may be an epicycle gear train,
such as a planetary gear system or other gear system, with a gear
reduction ratio of greater than about 2.3:1. It should be
understood, however, that the above parameters are only exemplary
of one embodiment of a geared architecture engine and that the
present disclosure is applicable to other gas turbine engines
including direct drive turbofans, land based turbine engines
utilized for power generation as well as turbine engines for use in
land based vehicles and sea going vessels.
[0041] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The
flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with
the engine at its best fuel consumption--also known as "bucket
cruise Thrust Specific Fuel Consumption (`TSFC`)" is the industry
standard parameter of lbm of fuel being burned divided by lbf of
thrust the engine produces at that minimum point. "Low fan pressure
ratio" is the pressure ratio across the fan blade alone, without a
Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as
disclosed herein according to one non-limiting embodiment is less
than about 1.45. "Low corrected fan tip speed" is the actual fan
tip speed in ft/sec divided by an industry standard temperature
correction of [(Tram .degree.R)/(518.7 .degree.R)].sup.0.5. The
"Low corrected fan tip speed" as disclosed herein according to one
non-limiting embodiment is less than about 1150 ft/second (350.5
meters/second).
[0042] The example gas turbine engine includes the fan 42 that
comprises in one non-limiting embodiment less than about twenty-six
(26) fan blades. In another non-limiting embodiment, the fan
section 22 includes less than about twenty (20) fan blades.
Moreover, in one disclosed embodiment the low pressure turbine 46
includes no more than about six (6) turbine rotors schematically
indicated at 34. In another non-limiting example embodiment, the
low pressure turbine 46 includes about three (3) turbine rotors. A
ratio between the number of fan blades 42 and the number of low
pressure turbine rotors is between about 3.3 and about 8.6. The
example low pressure turbine 46 provides the driving power to
rotate the fan section 22 and therefore the relationship between
the number of turbine rotors 34 in the low pressure turbine 46 and
the number of blades 42 in the fan section 22 disclose an example
gas turbine engine 20 with increased power transfer efficiency.
[0043] Referring to FIG. 2 with continued reference to FIG. 1, the
example combustor 56 includes an axially aft end 62 that is
adjacent to an axially forward face 72 of a first turbine vane
stage 64. The first turbine vane stage 64 includes an upper
platform 70 that defines the forward face 72 and has a radially
outward extending rib 92. The combustor 56 includes a rib 84 that
extends radially outward and is spaced apart from the end of the
combustor 56.
[0044] A conformal seal 76 is disposed between the rib 84 and the
forward face 72 on the radially outer surface 74 of the combustor
56. The conformal seal 76 extends axially aft from the rib 84 over
a radially extending gap 78 between the combustor 56 and the first
turbine vane stage 64.
[0045] Referring to FIG. 3 with continued reference to FIG. 2, the
first turbine vane stage 64 includes a plurality of turbine vanes
65 that extend between the upper platform 70 and a lower platform
75. Each vane 65 includes a leading edge 68 facing toward the
combustor 56. The leading edge 68 encounters the high-energy gas
flow generated in the combustor 56 and directs that gas flow into
the turbine section 28. The leading edge 68 of each vane 65 can
cause undesired distortions in gas flow that generate non-uniform
temperature variations within the gap 78. Bow wave flow phenomena
is one such flow distortion that may cause undesired discreet
temperature increases. Other flow disruptions that result in gas
flow entering the gap 78 may also result in undesired localized
temperature variations and also will benefit from this
disclosure.
[0046] The example turbine stage 64 includes a plurality of
doublets 66 that are arranged circumferentially about the engine
axis A. Each of the doublets 66 includes two vanes 65 with common
upper and lower platforms 70, 75. It is within the contemplation of
this disclosure to utilize other turbine vane stage configurations
with the disclosed seal 76.
[0047] Referring to FIG. 4 with continued reference to FIGS. 2 and
3, the disclosed example conformal seal 76 includes a first
plurality of cooling holes 80 that extend from a radially outer
surface 96 to a radially inner surface 98 that is in communication
with the gap 78. Each of the cooling holes 80 are disposed at an
angle 100 relative to the radially outer surface 96 such that
cooling air schematically shown at 102 exits into the gap 78. The
conformal seals 76 includes a wearing end portion 82 that wears
down during initial operation to provide a desired seal against the
axial face 72.
[0048] Referring to FIG. 5 with continued reference to FIGS. 2, 3
and 4, the axial face 72 includes a second plurality of cooling air
holes 88 that communicate cooling air into the gap 78. The cooling
air holes 88 are disposed at an angle 105 relative to the axial
face 72 to direct airflow radially inward into the gap 78.
[0049] The first and second plurality of cooling air holes 80, 88
are sized to provide a desired pressure and cooling airflow into
the gap 78. In one disclosed embodiment, the first plurality of
cooling air holes 80 are 0.025 inch (0.635 mm) in diameter and the
second plurality of cooling air holes 88 are 0.023 inch (0.5842 mm)
in diameter. According to another embodiment, the cooling air holes
may vary from between 0.015 inch (0.381 mm) and 0.080 inch (2.032
mm) in diameter. It should be understood that although an example
size of hole is disclosed by way of example, other sizes and
combinations of cooling hole structures are within the
contemplation of this disclosure.
[0050] Referring to FIGS. 6 and 7 with continued reference to FIGS.
2-5, the example conformal seal 76 includes a plurality of slots 86
arranged circumferentially about the engine axis A. Each of the
slots 86 is aligned with a corresponding leading edge 68 of the
vanes 65 within the first turbine vane section 64. The cooling
holes 80 open on the radial inner surface 98 of the seal 76 within
each of the plurality of slots 86. The cooling holes 80 communicate
cooling airflow to the gap 78 at a circumferential location that
corresponds with the leading edge 68 of each of the vanes 65.
[0051] The first and second plurality of cooling air holes 80 and
88 are grouped at the circumferential location that corresponds
with the leading edge 68. In one disclosed embodiment, each
grouping includes between 1 and 10 holes. In other disclosed
embodiment, each grouping of cooling air holes includes 8 holes.
While specific grouping counts are disclosed, other grouping counts
are within the contemplation of this disclosure.
[0052] The conformal seal 76 includes a tab 104 with a slot 106.
The slot 106 corresponds with slots 95 defined in rib 92 of the
vane stage 64 and slot 85 defined as part of the combustor rib 84.
An alignment member 94 extends through the slots 85, 106 and 95 to
align the slots 86 and the cooling holes 80 with the leading edge
68 of each vane 65.
[0053] Referring to FIG. 8 with continued reference to FIGS. 4, 5,
6 and 7, each of the slots 86 provides for communication of cooling
airflow 102 into the gap 78 in a location corresponding with the
leading edge 68 of each of the vanes 65. The cooling holes 80 are
grouped circumferentially about the circumference of the conformal
seal 76 to correspond with each of the slots 86.
[0054] The first and second plurality of cooling air holes 80 and
88 may both communicate air into each of the slots 86 defined
within the seal 76. The first and second plurality of cooling air
holes 80, 88 may also be incremented such that a first
circumferential position includes cooling air holes 80 from the
seal 76 and a second circumferential position includes cooling air
provide by the second plurality of cooling air holes 88. The
orientation of cooling air holes 80, 88 can be incremented, and
combined to provide cooling air flow through one or both the first
and second plurality of cooling air holes 80, 88 to provide desired
airflow required to maintain the gap 78 within desired temperature
ranges.
[0055] Referring to FIG. 9A, a set of the plurality of slots 86
indicated at 110 receives cooling air from the first plurality of
cooling air holes 80 and a second set of the plurality of slots 86
indicated at 112 receives cooling air from the second plurality of
cooling air holes 88. Referring to FIG. 9B, another example
disclosed embodiment is schematically illustrated to show another
circumferential orientation of the first set of slots 110 and the
second set of slots 112. It should be appreciated that the sets of
slots need not be symmetric about the engine axis and can include
different combination of cooling air holes communicating cooling
air into the plurality of slots 86.
[0056] Cooling airflow 102 is communicated through the conformal
seal 76 and vane stage 64 into the gap 78 at the specific
circumferential location that corresponds with the leading edge 68
of each of the vanes 65. Accordingly, the example conformal seal 76
provides a seal between the end of the combustor and the first
turbine stage while also providing directed cooling airflow to
prevent or substantially limit hot gas flow into the gap 78.
[0057] Although an example embodiment has been disclosed, a worker
of ordinary skill in this art would recognize that certain
modifications would come within the scope of this disclosure. For
that reason, the following claims should be studied to determine
the scope and content of this disclosure.
* * * * *